US6546731B2 - Combustion chamber for a gas turbine engine - Google Patents

Combustion chamber for a gas turbine engine Download PDF

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Publication number
US6546731B2
US6546731B2 US09/726,194 US72619400A US6546731B2 US 6546731 B2 US6546731 B2 US 6546731B2 US 72619400 A US72619400 A US 72619400A US 6546731 B2 US6546731 B2 US 6546731B2
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Prior art keywords
holes
wall
effusion
impingement
hole
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US09/726,194
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US20010004835A1 (en
Inventor
Hisham Salman Alkabie
Robin Thomas David McMillan
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Siemens AG
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Alstom Power UK Holdings Ltd
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Assigned to ABB ALSTOM POWER UK LTD. reassignment ABB ALSTOM POWER UK LTD. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MCMILLAN, ROBIN THOMAS DAVID, ALKABIE, HISHAM SALMAN
Publication of US20010004835A1 publication Critical patent/US20010004835A1/en
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Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ALSTOM POWER UK HOLDINGS FORMERLY ALSTOM POWER UK LTD. FORMERLY ABB ALSTOM POWER UK LTD.
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • This invention relates to gas turbine engines, and in particular to cooling of combustion chamber walls in such engines.
  • combustion chambers in gas turbine engines are subject to very high temperatures in use and, as efforts are made to increase engine efficiency, higher operating temperatures become desirable.
  • higher operating temperatures become desirable.
  • the ability of the combustion chamber walls to withstand higher temperatures becomes a limiting factor in engine development.
  • New wall materials to withstand higher temperatures are constantly being developed, but there is usually some cost or functional penalty involved.
  • metal alloys become more exotic, they tend to be more expensive, both in the materials required and in the complexity of manufacture.
  • Ceramic materials on the other hand, while being able to withstand high temperatures, tend to exhibit low mechanical strength.
  • the combustion chamber is formed with twin walls spaced apart from each other by a small distance.
  • Compressed air from the engine compressor surrounds the combustion chambers within the engine casing, and holes formed in the outer wall of the twin walls of the chamber allow air to impinge on the inner wall, creating a first cooling effect.
  • Such holes are normally referred to as impingement holes.
  • the air in the space between the walls is then admitted to the combustion chamber through a series of smaller holes, normally referred to as effusion holes, through the inner wall which are arranged to aid laminar flow of the cooling air in a film over the inner surface of the inner wall, cooling it and providing a protective layer from the combustion gases in the chamber.
  • effusion holes through the inner wall which are arranged to aid laminar flow of the cooling air in a film over the inner surface of the inner wall, cooling it and providing a protective layer from the combustion gases in the chamber.
  • a combustion chamber for a gas turbine engine having:
  • the outer wall having a plurality of impingement cooling holes therethrough, whereby, during operation of the engine, compressed air surrounding the chamber can pass through the impingement holes to impinge on the inner wall,
  • the inner wall having a plurality of effusion holes therethrough, whereby air can effuse from the cavity between the inner and outer walls into the combustion chamber, there being a greater number of effusion holes than impingement holes,
  • effusion holes are arranged in groups, each group comprising a plurality of effusion holes substantially equally spaced apart from each other around a central effusion hole, each group of effusion holes having an impingement hole located in the outer wall such that air passing through the impingement hole impinges on the inner wall at a predetermined position relative to the central effusion hole within a boundary defined by the group of diffusion holes.
  • the effusion holes are arranged in groups of seven, comprising six effusion holes substantially equally spaced around a central seventh effusion hole.
  • the predetermined position of the impingement hole relative to the central effusion hole is preferably such that air passing through the impingement hole impinges on the inner wall closer to the central effusion hole than to the other effusion holes and is in alignment with the central effusion hole along the direction of combustion gas flow in the chamber.
  • each impingement hole may be located upstream or downstream of the central effusion hole in the group, but is more preferably arranged downstream of the central effusion hole such that the centerline of the impingement hole is spaced from the centerline of the central effusion hole by a distance at least equal to the diameter of the impingement hole.
  • the groups are suitably arranged in rows extending circumferentially of the chamber.
  • each group may be spaced from the next in the row by a distance substantially equal to the spacing between adjacent holes in a group and the groups in any one row may be displaced circumferentially from those in the or each adjacent row by a distance substantially equal to half the distance between the central holes in adjacent groups in a row.
  • the longitudinal spacing between the rows may be such that the distance between two adjacent effusion holes which belong to different groups in adjacent rows is the same as the distance between two adjacent holes in the same group of effusion holes.
  • additional effusion holes are provided centrally of each set of six holes defined between two adjacent groups in one row and the displaced adjacent group in the next row.
  • the relative sizes and numbers of the impingement holes and the effusion holes are preferably such that, during operation of the engine, the pressure differential across the outer wall is at least twice the pressure differential across the inner wall; for example, approximately 70% of the total pressure drop across the outer and inner walls may occur across the outer wall and the remainder across the inner wall.
  • the combustion chamber wall temperature during operation of the engine is significantly lower using the arrangement of the invention than is achieved with known cooling arrangements.
  • Benefits are gained from the enhanced film cooling not only in the combustion chamber can, but also into the transition duct which leads from the can into the turbine inlet.
  • the enhanced cooling extends the life of the combustion chamber can and its transition duct, especially when combustion temperatures are increased to improve combustion efficiency.
  • FIG. 1 is a diagrammatic sectional view of a combustion chamber
  • FIG. 2 is an enlarged partial view of the wall of the combustion chamber within box A in FIG. 1;
  • FIG. 3 is an enlarged plan diagram showing the arrangement of cooling holes in a single group of such holes
  • FIG. 4 is a view similar to FIG. 3 but on a reduced scale and showing the relationship between adjacent groups of cooling holes in accordance with one embodiment of the invention.
  • FIG. 5 is a corresponding view to that of FIG. 4, but showing an alternative embodiment of the invention.
  • the combustion chamber can 1 has a conventional inlet or upstream end 10 for fuel and combustion air, and a discharge or downstream end 12 , the flow of the combustion air and combustion gases through the chamber being indicated by arrows B and D respectively.
  • Downstream of the inlet end 10 the can is generally cylindrical about its longitudinal axis L-L and has twin walls 2 , 4 spaced apart by a small distance in conventional manner to provide a cooling air space cavity 13 between them:
  • the structure of the twin walls may be seen more clearly from FIG. 2, with the outer wall 2 being provided with impingement holes 3 therethrough, while the inner wall 4 has effusion holes 5 therethrough. Although the impingement holes are shown in FIG.
  • the effusion holes are conveniently formed by laser drilling. It will be seen that the impingement holes are arranged such that during operation of the engine, compressed air C from the space within the engine casing surrounding the combustion chamber 1 flows into the cavity 13 between the walls 2 and 4 and impinges directly on the hot inner wall 4 at a position offset from the positions of the effusion holes 5 so that an initial cooling effect on inner wall 4 is achieved by the impingement.
  • the effusion holes 5 are arranged in polygonal groups, each group comprising a number of effusion holes 5 a substantially equally spaced apart from each other around a central effusion hole 5 b .
  • Each group of effusion holes is associated with a respective impingement hole 3 which is located in the outer wall 2 such that air passing through the impingement hole impinges on the inner wall 4 at a predetermined position 14 relative to the central effusion hole. This center of impingement 14 is within the polygonal boundary defined by the diffusion holes 5 a.
  • air passing through the impingement holes 3 impinges on the inner wall 4 closer to the central effusion hole 5 b than to the other effusion holes 5 a , the center of impingement 14 being in alignment with the central effusion hole 5 b along the direction D of combustion gas flow in the chamber, and preferably downstream of hole 5 b.
  • the effusion holes 5 are arranged in the inner wall 4 in groups of seven as shown, with each of six holes 5 a defining with the next adjacent hole an equal side of a hexagon, the seventh effusion hole 5 b being at the center of the hexagon.
  • the impingement hole 3 in the outer wall 2 associated with the group is positioned downstream of the central effusion hole 5 b such that the horizontal distance d between the centerline of the central hole 5 b and the centerline of the impingement hole 3 is at least equal to the diameter of the impingement hole. It will be seen that the impingement holes 3 have a significantly greater diameter than the effusion holes, although the number of effusion holes is substantially greater than the number of impingement holes.
  • the relative sizes and numbers of the two types of hole are designed to ensure that the pressure differential across the outer wall 2 is at least twice the pressure differential across the inner wall 4 .
  • FIG. 4 One exemplary arrangement of the groups of effusion holes is shown in FIG. 4 .
  • the groups G 1 , G 2 , etc. are arranged in parallel rows R 1 , R 2 , etc., extending circumferentially around the can.
  • each group G 1 is spaced from the next group G 2 in the row by a distance S, which as shown is also the spacing between adjacent holes in a group along each side of the hexagon in which they are arranged.
  • the groups in one row R 1 are offset circumferentially from those in the next adjacent row R 2 by half the distance X between the adjacent central holes 5 b 1 , 5 b 2 .
  • the longitudinal spacing between the rows is such that the distance between two adjacent effusion holes which belong to different groups in adjacent rows is the same as the distance between two adjacent holes in the same group.
  • effusion hole 5 a 1 in group G 1 of row R 1 and an adjacent effusion hole 5 a 2 of another group in the adjacent row R 2 the distance between them is S.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US09/726,194 1999-12-01 2000-11-29 Combustion chamber for a gas turbine engine Expired - Lifetime US6546731B2 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
GB9928242A GB2356924A (en) 1999-12-01 1999-12-01 Cooling wall structure for combustor
GB9928242 1999-12-01
GB9928242.8 1999-12-01

Publications (2)

Publication Number Publication Date
US20010004835A1 US20010004835A1 (en) 2001-06-28
US6546731B2 true US6546731B2 (en) 2003-04-15

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US (1) US6546731B2 (de)
EP (1) EP1104871B1 (de)
JP (1) JP4554802B2 (de)
DE (1) DE60012289T2 (de)
ES (1) ES2223410T3 (de)
GB (1) GB2356924A (de)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20030182943A1 (en) * 2002-04-02 2003-10-02 Miklos Gerendas Combustion chamber of gas turbine with starter film cooling
US20030200752A1 (en) * 2002-04-29 2003-10-30 Moertle George Eric Multihole patch for combustor liner of a gas turbine engine
US6868675B1 (en) 2004-01-09 2005-03-22 Honeywell International Inc. Apparatus and method for controlling combustor liner carbon formation
US20050241321A1 (en) * 2004-04-30 2005-11-03 Martling Vincent C Transition duct apparatus having reduced pressure loss
US20050241316A1 (en) * 2004-04-28 2005-11-03 Honeywell International Inc. Uniform effusion cooling method for a can combustion chamber
US20070028595A1 (en) * 2005-07-25 2007-02-08 Mongia Hukam C High pressure gas turbine engine having reduced emissions
US20070271925A1 (en) * 2006-05-26 2007-11-29 Pratt & Whitney Canada Corp. Combustor with improved swirl
US20070271926A1 (en) * 2006-05-26 2007-11-29 Pratt & Whitney Canada Corp. Noise reducing combustor
US20080127651A1 (en) * 2006-11-30 2008-06-05 Honeywell International, Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20100000200A1 (en) * 2008-07-03 2010-01-07 Smith Craig F Impingement cooling device
US20100170257A1 (en) * 2009-01-08 2010-07-08 General Electric Company Cooling a one-piece can combustor and related method
US20100257863A1 (en) * 2009-04-13 2010-10-14 General Electric Company Combined convection/effusion cooled one-piece can combustor
US20100272953A1 (en) * 2009-04-28 2010-10-28 Honeywell International Inc. Cooled hybrid structure for gas turbine engine and method for the fabrication thereof
US20110016874A1 (en) * 2009-07-22 2011-01-27 Rolls-Royce Plc Cooling Arrangement for a Combustion Chamber
US20120255308A1 (en) * 2011-04-06 2012-10-11 Rolls-Royce Plc Cooled double walled article
US8438856B2 (en) 2009-03-02 2013-05-14 General Electric Company Effusion cooled one-piece can combustor
US8834154B2 (en) * 2012-11-28 2014-09-16 Mitsubishi Heavy Industries, Ltd. Transition piece of combustor, and gas turbine having the same
US9052111B2 (en) 2012-06-22 2015-06-09 United Technologies Corporation Turbine engine combustor wall with non-uniform distribution of effusion apertures
US9157328B2 (en) 2010-12-24 2015-10-13 Rolls-Royce North American Technologies, Inc. Cooled gas turbine engine component
US20160010862A1 (en) * 2014-07-14 2016-01-14 Rolls-Royce Plc Annular combustion chamber wall arrangement
US20220162955A1 (en) * 2019-03-04 2022-05-26 Rolls-Royce Deutschland Ltd & Co Kg Method for manufacturing an engine component with a cooling duct arrangement and engine component

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GB2361303B (en) * 2000-04-14 2004-10-20 Rolls Royce Plc Wall structure for a gas turbine engine combustor
US7475853B2 (en) * 2002-06-21 2009-01-13 Darko Segota Method and system for regulating external fluid flow over an object's surface, and particularly a wing and diffuser
US7296411B2 (en) * 2002-06-21 2007-11-20 Darko Segota Method and system for regulating internal fluid flow within an enclosed or semi-enclosed environment
US20050098685A1 (en) * 2002-06-21 2005-05-12 Darko Segota Method and system for regulating pressure and optimizing fluid flow about a fuselage similar body
US7048505B2 (en) * 2002-06-21 2006-05-23 Darko Segota Method and system for regulating fluid flow over an airfoil or a hydrofoil
US6964170B2 (en) * 2003-04-28 2005-11-15 Pratt & Whitney Canada Corp. Noise reducing combustor
US7036316B2 (en) * 2003-10-17 2006-05-02 General Electric Company Methods and apparatus for cooling turbine engine combustor exit temperatures
US7531048B2 (en) * 2004-10-19 2009-05-12 Honeywell International Inc. On-wing combustor cleaning using direct insertion nozzle, wash agent, and procedure
EP1650503A1 (de) * 2004-10-25 2006-04-26 Siemens Aktiengesellschaft Verfahren zur Kühlung eines Hitzeschildelements und Hitzeschildelement
US7827801B2 (en) * 2006-02-09 2010-11-09 Siemens Energy, Inc. Gas turbine engine transitions comprising closed cooled transition cooling channels
DE102006042124B4 (de) * 2006-09-07 2010-04-22 Man Turbo Ag Gasturbinenbrennkammer
JP5296320B2 (ja) * 2007-01-30 2013-09-25 ゼネラル・エレクトリック・カンパニイ 逆流噴射機構を有するシステム及び燃料及び空気を噴射する方法
US7886517B2 (en) * 2007-05-09 2011-02-15 Siemens Energy, Inc. Impingement jets coupled to cooling channels for transition cooling
US7617684B2 (en) * 2007-11-13 2009-11-17 Opra Technologies B.V. Impingement cooled can combustor
US20100037620A1 (en) * 2008-08-15 2010-02-18 General Electric Company, Schenectady Impingement and effusion cooled combustor component
US8590314B2 (en) * 2010-04-09 2013-11-26 General Electric Company Combustor liner helical cooling apparatus
US8647053B2 (en) 2010-08-09 2014-02-11 Siemens Energy, Inc. Cooling arrangement for a turbine component
JP5821550B2 (ja) * 2011-11-10 2015-11-24 株式会社Ihi 燃焼器ライナ
EP2644995A1 (de) 2012-03-27 2013-10-02 Siemens Aktiengesellschaft Verbesserte Lochanordnung von Auskleidungen einer Brennkammer eines Gasturbinenmotors mit niedriger Verbrennungsdynamik und niedrigen Emissionen
DE102012025375A1 (de) * 2012-12-27 2014-07-17 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Anordnung von Prallkühllöchern und Effusionslöchern in einer Brennkammerwand einer Gasturbine
EP3077641B1 (de) * 2013-12-06 2020-02-12 United Technologies Corporation Kühlung einer zünderdurchführung einer brennkammerwand
US10094564B2 (en) * 2015-04-17 2018-10-09 Pratt & Whitney Canada Corp. Combustor dilution hole cooling system
GB201518345D0 (en) * 2015-10-16 2015-12-02 Rolls Royce Combustor for a gas turbine engine
DE102016219424A1 (de) * 2016-10-06 2018-04-12 Rolls-Royce Deutschland Ltd & Co Kg Brennkammeranordnung einer Gasturbine sowie Fluggasturbine
US10697635B2 (en) * 2017-03-20 2020-06-30 Raytheon Technologies Corporation Impingement cooled components having integral thermal transfer features
US11028705B2 (en) * 2018-03-16 2021-06-08 Doosan Heavy Industries Construction Co., Ltd. Transition piece having cooling rings
KR102593506B1 (ko) * 2018-09-11 2023-10-24 한화에어로스페이스 주식회사 가스 터빈 장치의 케이스 구조체

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Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20030182943A1 (en) * 2002-04-02 2003-10-02 Miklos Gerendas Combustion chamber of gas turbine with starter film cooling
US7124588B2 (en) * 2002-04-02 2006-10-24 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber of gas turbine with starter film cooling
US20030200752A1 (en) * 2002-04-29 2003-10-30 Moertle George Eric Multihole patch for combustor liner of a gas turbine engine
US7086232B2 (en) * 2002-04-29 2006-08-08 General Electric Company Multihole patch for combustor liner of a gas turbine engine
US6868675B1 (en) 2004-01-09 2005-03-22 Honeywell International Inc. Apparatus and method for controlling combustor liner carbon formation
US20060207095A1 (en) * 2004-01-09 2006-09-21 Honeywell International Inc. Method for controlling carbon formation on repaired combustor liners
US7124487B2 (en) 2004-01-09 2006-10-24 Honeywell International, Inc. Method for controlling carbon formation on repaired combustor liners
US20050241316A1 (en) * 2004-04-28 2005-11-03 Honeywell International Inc. Uniform effusion cooling method for a can combustion chamber
US20050241321A1 (en) * 2004-04-30 2005-11-03 Martling Vincent C Transition duct apparatus having reduced pressure loss
US7137241B2 (en) 2004-04-30 2006-11-21 Power Systems Mfg, Llc Transition duct apparatus having reduced pressure loss
US20070028595A1 (en) * 2005-07-25 2007-02-08 Mongia Hukam C High pressure gas turbine engine having reduced emissions
US20070271925A1 (en) * 2006-05-26 2007-11-29 Pratt & Whitney Canada Corp. Combustor with improved swirl
US20070271926A1 (en) * 2006-05-26 2007-11-29 Pratt & Whitney Canada Corp. Noise reducing combustor
US7628020B2 (en) 2006-05-26 2009-12-08 Pratt & Whitney Canada Cororation Combustor with improved swirl
US7856830B2 (en) 2006-05-26 2010-12-28 Pratt & Whitney Canada Corp. Noise reducing combustor
US20080127651A1 (en) * 2006-11-30 2008-06-05 Honeywell International, Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US7926284B2 (en) 2006-11-30 2011-04-19 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20100000200A1 (en) * 2008-07-03 2010-01-07 Smith Craig F Impingement cooling device
US9046269B2 (en) 2008-07-03 2015-06-02 Pw Power Systems, Inc. Impingement cooling device
US20100170257A1 (en) * 2009-01-08 2010-07-08 General Electric Company Cooling a one-piece can combustor and related method
US8438856B2 (en) 2009-03-02 2013-05-14 General Electric Company Effusion cooled one-piece can combustor
US20100257863A1 (en) * 2009-04-13 2010-10-14 General Electric Company Combined convection/effusion cooled one-piece can combustor
US20100272953A1 (en) * 2009-04-28 2010-10-28 Honeywell International Inc. Cooled hybrid structure for gas turbine engine and method for the fabrication thereof
EP2292977A2 (de) 2009-07-22 2011-03-09 Rolls-Royce plc Kühlanordnung für eine Brennkammer
US8794961B2 (en) * 2009-07-22 2014-08-05 Rolls-Royce, Plc Cooling arrangement for a combustion chamber
EP2292977A3 (de) * 2009-07-22 2016-05-18 Rolls-Royce plc Kühlanordnung für eine Brennkammer
US20110016874A1 (en) * 2009-07-22 2011-01-27 Rolls-Royce Plc Cooling Arrangement for a Combustion Chamber
US9157328B2 (en) 2010-12-24 2015-10-13 Rolls-Royce North American Technologies, Inc. Cooled gas turbine engine component
US9010124B2 (en) * 2011-04-06 2015-04-21 Rolls-Royce Plc Cooled double walled article
US20120255308A1 (en) * 2011-04-06 2012-10-11 Rolls-Royce Plc Cooled double walled article
US9052111B2 (en) 2012-06-22 2015-06-09 United Technologies Corporation Turbine engine combustor wall with non-uniform distribution of effusion apertures
US8834154B2 (en) * 2012-11-28 2014-09-16 Mitsubishi Heavy Industries, Ltd. Transition piece of combustor, and gas turbine having the same
US20160010862A1 (en) * 2014-07-14 2016-01-14 Rolls-Royce Plc Annular combustion chamber wall arrangement
US10563866B2 (en) * 2014-07-14 2020-02-18 Rolls-Royce Plc Annular combustion chamber wall arrangement
US20220162955A1 (en) * 2019-03-04 2022-05-26 Rolls-Royce Deutschland Ltd & Co Kg Method for manufacturing an engine component with a cooling duct arrangement and engine component
US11939889B2 (en) * 2019-03-04 2024-03-26 Rolls-Royce Deutschland Ltd & Co Kg Method for manufacturing an engine component with a cooling duct arrangement and engine component

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ES2223410T3 (es) 2005-03-01
US20010004835A1 (en) 2001-06-28
EP1104871B1 (de) 2004-07-21
GB9928242D0 (en) 2000-01-26
JP2001227359A (ja) 2001-08-24
DE60012289T2 (de) 2005-07-28
DE60012289D1 (de) 2004-08-26
JP4554802B2 (ja) 2010-09-29
EP1104871A1 (de) 2001-06-06
GB2356924A (en) 2001-06-06

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